The fuselage, wings, and empennage of an aircraft typically include stringers that are coupled to skin that forms the smooth aerodynamic outer surfaces of the fuselage, wings, and empennage. The stringers and skin cooperate to provide flexural and torsional stiffness to these sections of the aircraft. Traditionally, the fuselage, wings, and empennage surfaces and the associated stringers are fabricated from metal, such as aluminum, steel, or titanium. The stringer may include a web portion, such as a planar wall, that is generally oriented in a direction approximately perpendicular to the skin and extends in a generally lengthwise direction along the fuselage and empennage and in a generally spanwise direction along the wing so that the web portion provides resistance to bending. A flange portion may be positioned on one or both of the longitudinal edges of the web portion to provide increased rigidity and support to the stringer. The flange portion along one of the longitudinal edges of the web portion can also be used as an attachment surface for attaching the stringer to the skin.
Fiber reinforced composite materials are widely used in a variety of commercial and military aircraft products as a substitute for metals, particularly in applications where relatively low weight and high mechanical strength are desired. The material is generally comprised of a network of reinforcing fibers that are arranged in layers or plies. The layers include a resin matrix that substantially wets the reinforcing fibers and that is cured to form an intimate bond between the resin and the reinforcing fibers. The composite material may be formed into a structural component by a variety of known forming methods, such as extrusion, vacuum bagging, autoclaving, and/or the like.
As the skins and stringers for various sections of aircrafts transition from metallic materials to fiber reinforced composite materials, multiple issues have arisen. In a current fabrication process, a fiber reinforced composite skin is formed by stacking layers together that contain reinforcing fibers in a resin matrix. Typically, some of the layers are staggered relative to each other so that the stack conforms to a desired contoured or tapered geometry. The stacked layers are heated and pressurized to cure the polymeric resin matrix and form a precured skin. Hard tooling, e.g., metallic tooling or metallic mold/die, containing an uncured fiber reinforced composite material that is shaped or preformed into a stringer is positioned along the precured skin. Pressure and heat are applied to cure the stringer preform using the hard tooling to form a fiber reinforced composite stringer that is attached to the precured skin. Unfortunately, defects often occur along the interface between the precured skin and the fiber reinforced composite stringer. In particular, the precured skin typically has a contoured outer surface that includes small steps or drop-offs that are formed by the staggered, stacked layers of fibers and polymer resin. As such, it is difficult to match and position the hard tooling to continuously follow the outer surface of the precured skin and the hard tooling will often bridge across these sections of the precured skin forming under compressed areas, e.g., voids, and over compressed areas, e.g., resin poor areas, at the interface between the precured skin and the fiber reinforced composite stringer. These under and over compressed areas can reduce the load transfer efficacy between the precured skin and the fiber reinforced composite stringer, thereby reducing the rigidity and support provided by the fiber reinforced composite stringer. Moreover, fiber reinforced composite stringers formed by this fabrication process or similar fabrication processes typically have a flange portion positioned along only one of the longitudinal edges of the web portion, particularly in the runout or end portions of the stringer, because it is difficult to remove the hard tooling from the fiber reinforced composite stringer after curing. As such, the rigidity and support of the fiber reinforced composite stringer is further compromised.
Accordingly, it is desirable to provide reinforced composite structures for an aircraft including fiber reinforced composite stringers and, optionally, fiber reinforced composite skins affixed to the fiber reinforced composite stringers that provide improved flexural and torsional stiffness, and methods for making such reinforced composite structures. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description and the appended claims, taken in conjunction with the accompanying drawings and this background.